1. Field of the Invention
The invention is generally related to solar thermal rockets and more particularly to the use of thermal energy storage modules in solar thermal rockets.
2. General Background
Solar thermal rockets were first proposed in 1954 as a way to provide greater specific impulse than chemical rockets. Solar thermal rockets use the sun""s energy to heat a propellant (typically hydrogen) to extremely high temperatures and then expel the gas through a nozzle to produce thrust. The high temperature and low molecular weight of the propellant combine to produce a specific impulse of two to four times that of a chemical rocket. Generally, solar thermal rockets have been of a xe2x80x9cdirect gainxe2x80x9d design, in which the propellant is heated directly by incident concentrated sunlight during a propulsive burn. Direct gain engines offer the capability to operate at very high temperatures (theoretically greater than 3000 K) resulting in a very high specific impulse (theoretically greater than 950 seconds (ideal) for hydrogen). Material limitations typically limit the realized ideal specific impulse to less than 900 seconds. The disadvantage of direct gain systems is that they require very large, highly efficient primary solar concentrators (either alone or in combination with secondary concentrators) to provide the high power required to raise the propellant temperature to desired operating levels to yield thrust levels of interest. To date, these primary concentrators do not exist. In addition, direct gain systems must continually point their concentrators accurately toward the sun while thrusting. This places a premium on the overall system pointing and tracking requirements as well as the thrust vector control of the engine in order to ensure that the thrust is continuously in the direction required.
Stored thermal energy systems collect and store the incident solar energy over a relatively long period and then transfer the energy to the propellant during a short propulsive burn. The thermal energy storage design solves the primary concentrator problem by using existing smaller primary concentrators to collect and store solar energy over one or more orbital periods and then using the stored energy to heat the propellant over a short pulse burn. Several such heat-up and burn cycles (charge/discharge cycles) are performed to move the satellite to its destination. The longer the charge period of each cycle, the smaller the primary concentrator can be. This approach thus enables the use of existing primary concentrator technologies to develop an operational system, and permits higher thrust levels since thrust is decoupled from primary concentrator size. In addition, thrust operations can be performed without the need to maintain solar pointing during the propulsive maneuver, thereby simplifying the pointing and tracking hardware and software.
Stored thermal energy systems typically use either phase change material (which stores energy in the latent heat associated with changing from solid to liquid or liquid to gas) or solid, lightweight, high thermal capacity materials (which store the energy as sensible heat). Phase change systems typically operate at lower temperatures while sensible heat systems can operate at very high temperatures. A number of drawbacks exist for these systems however. The major disadvantage in high temperatures systems is that the energy storage materials (typically rhenium-encapsulated graphite or tungsten encapsulated boron nitride) have temperature limitations below direct gain systems. Dual material systems are needed since the high specific heat materials used as the thermal storage media tend to have high vapor pressures and react chemically with the hydrogen propellant. Coating of the storage media with high temperature metals or ceramics is required to provide long life. Material stability of these combined systems is challenged at the temperatures direct gain systems can operate. In order to offset this deficiency, a thermal energy storage design has to operate at higher thrust levels to achieve the same delivered impulse. Past designs have sought to apply the thin protective coatings directly to the storage material and rely on the storage material to serve as the structural member. In general, this has been very difficult to achieve on a consistent basis. Because the coating serves as a pressure boundary, it must be hermetically tight. If any region of the coating is found to leak, the whole surface must be coated further until any leakage is eliminated. Hence, manufacturing control and reliability are nearly impossible to achieve. Furthermore, failure of the coating or problems during the coating process can render the piece useless. An additional detriment is that since the storage material serves as the structural element, any change in size dictates a whole new design, fabrication, and design qualification process.
This invention addresses the above need. What is provided is a modular solar thermal rocket that receives and absorbs solar energy and then acts as a heat exchanger to provide propulsive thrust. This receiver/absorber/exchanger (RAX) is comprised of several thermal energy storage modules. The thermal energy storage modules (TEM) receive and store solar energy via thermal energy storage (TES) elements provided in each module. The solar energy from the primary concentrator is focused into a secondary concentrator which further focuses the sunlight into the cavity which is formed by the arrayed thermal energy storage modules. A preheater is positioned adjacent to the secondary concentrator and is in fluid communication with a propellant supply and a common header which feeds one end of each of the thermal energy storage modules. A propulsion nozzle is in fluid communication with the opposite end of the thermal energy storage modules. Stored propellant is directed through the preheater and the thermal energy storage modules where it is heated to a high temperature. The propellant is then directed to the propulsion nozzle where it is exhausted into space to provide propulsive thrust.